The present invention relates generally to gas turbine engines, and, more specifically, to aircraft turbofan engines.
Gas turbine engines have evolved into many configurations for powering various forms of commercial and military aircraft. The typical turbofan engine includes in serial flow communication a fan, compressor, combustor, high pressure turbine (HPT), and low pressure turbine (LPT).
Air enters the engine and is pressurized by the fan and compressor and mixed with fuel in the combustor for generating hot combustion gases. Energy is extracted from the combustion gases in the HPT which powers the compressor through an interconnecting shaft. Additional energy is extracted from the combustion gases in the LPT which powers the fan through a second shaft.
The fan is typically disposed inside a fan nacelle that defines a substantially annular bypass duct around the cowl which surrounds the core engine. Air pressurized by the fan blades is split radially with an inner portion being channeled through the compressor of the core engine, and an outer portion being channeled through the bypass duct, and therefore bypassing the core engine. Propulsion thrust is generated by the pressurized fan air bypassing the core engine as well as by the hot combustion gases discharged from the core engine.
Turbofan engines may be low or high bypass depending upon the amount of fan air bypassing the core engine. Modern turbofan aircraft engines powering commercial aircraft in flight are typically high bypass engines with relatively large, single stage fan blades mounted inside the nacelle and powered by a multistage LPT. The HPT may have a single stage or multiple stages therein and cooperates with the multiple stages of the LPT for maximizing energy extraction from the combustion gases to power the fan and compressor.
The compressor in a modern turbofan engine is typically a multistage axial high pressure compressor directly driven by the rotor or shaft of the HPT. And in some configurations, a multistage, axial booster or low pressure compressor is disposed between the fan and high pressure compressor and joined to the fan shaft or rotor powered by the LPT.
The compressors and turbines have various stages or rows of rotor blades extending radially outwardly from supporting rotor spools or disks joined together by the corresponding rotors or shafts. Each stage or row of rotor blades typically cooperates with an upstream row or stage of stator vanes.
Stator vanes and rotor blades have corresponding airfoil configurations which cooperate for pressurizing the air in the compressor and expanding the combustion gases in the turbines for extracting energy therefrom. Each airfoil has a generally concave pressure side and an opposite, generally convex suction side extending radially in span between axially opposite leading and trailing edges.
The nominal curvature of the airfoil is represented by the camber line extending between the leading and trailing edges. And, the concave pressure side and convex suction side are specifically configured for providing the desired pressure distributions thereover for maximizing efficiency of air compression in the compressor and gas expansion in the turbines.
The rotors of the HPT and LPT typically rotate in the same direction, or co-rotate, and the angular or twist orientation of the vanes and blades in the compressor and turbines typically alternate between the airfoil rows as the flow streams are turned in their tortuous path through the engine.
Each vane and blade row has a corresponding total number of airfoils therein required for efficiently turning the flow streams under the aerodynamic loading therefrom. Each row typically has a substantial number or multitude of airfoils around the circumference thereof dictated by the aerodynamic loading requirements of each stage and the turning or swirling of the flow streams axially therethrough.
For example, a single stage high pressure (HP) turbine typically has a substantial amount of exit swirl of the combustion gases, for example about 25 degrees. Correspondingly, the first stage low pressure (LP) turbine nozzle has vanes with substantial curvature or camber for efficiently turning the high swirl discharge flow from the HPT.
In a two stage HPT, the second stage HP blades typically have corresponding camber and angular orientation or twist relative to the axial centerline axis of the engine for effecting nearly zero swirl at the exit of the HPT. Correspondingly, the first stage LP nozzle vanes will have suitable camber and twist for efficiently channeling the combustion gases to the first stage LP blades.
Modern turbofan engines presently used for powering commercial aircraft in flight enjoy high operating efficiency due to the many advancements in design of the various components thereof over many years of development and commercial use in service. Since the engines power aircraft in flight, the size and weight of the engines themselves are ever paramount design objectives along with maximum efficiency of operation. The cost of jet fuel continually increases, and the need to further maximize efficiency of turbofan engines and reduce fuel consumption becomes ever more challenging in modern aircraft engine design.
Accordingly, it is desired to provide a turbofan aircraft engine having further improvement in efficiency in the turbine stages thereof.